Field of the Invention
The present invention relates to gas turbine engine combustors and more specifically to an improved combustor liner cooling system for achieving high combustion operating temperatures and efficiencies.
Conventional combustor cooling systems typically employ a set pattern of small diameter cooling holes drilled at an angle through the liner thickness. Cooling air passes through the holes, convectively cooling across the hole surfaces, and then exits into the main combustion gas stream as film, further enhancing the cooling effectiveness of the system.
As combustor and coolant temperature requirements increase in higher performance engines, the amount of film required to cool a metallic liner material also increases. Since the film does not contribute to burning in the combustor, the arrangement reduces the level of combustion temperature rise within the burner and stoichiometric temperatures are not achieved. With less air available for mixing with the fuel, there are reductions in combustor efficiency levels and engine performance.
Due to the temperature limitations of metal liners, composite or non-metallic liner materials having high temperature/strength capabilities relative to metals are being investigated. Materials such as carbon/carbon or ceramic matrix composite are strong candidates. However, a major disadvantage of carbon/carbon material is that if small cooling holes are drilled through the thickness, the carbon fibers will oxidize, resulting in ash with zero fiber strength.
Alternate techniques for cooling are convective cooling the backside of the combustor liner using air convection or impingement cooling. Backside convective cooling results in a hot surface temperature exceeding the material strength and temperature capabilities, and inadequate pressure to inject the air within the combustion zone area due to the high pressure drop assaciated with maintaining adequate air velocity for high convection heat transfer. Impingement cooling results in surface temperatures that are within the acceptable limits for carbon/carbon material, while also allowing adequate pressure to inject the cooling air into the flow stream. In this type of a system, the spent cooling air exits through holes in the liner. However, for the reasons stated above, the use of holes in a carbon/carbon material liner has drawbacks. Without an available exit, the spent impingement cooling air wi11 create a cross-flow condition which could reduce the cooling effectiveness of the system. The cooling effectiveness reductions may be large enough to cause an increase in liner temperatures which would exceed the material's temperature and strength capabilities.
U.S. Pat. No. 4,567,730 to Scott discloses a shielded combustor having liners made of non-metallic material such as ceramic or carbon/carbon capable of withstanding elevated temperatures. A plurality of cooling air apertures disposed in an outer shell channel high-speed jets of impingement cooling air upon the outer surface of the liners. A portion of the cooling air may flow through an optional dilution aperture in the liners into the combustion zone, and another portion is discharged downstream as film cooling.
Another prior art combustor cooling system is disclosed in U.S. Pat. No. 4,916,906 to Vogt. Method and apparatus are disclosed for providing breach cooling of an imperforate wall combustor liner. The breach cooling is effected by structure for channeling a cooling fluid such as a jet toward an outer surface of the imperforate wall, with the jet having sufficient momentum to breach a boundary layer of the cooling fluid which forms over the wall outer surface for more effective cooling. In an exemplary embodiment, the breach-cooled wall is an upstream portion of the gas turbine engine combustor, and the inner surface of the combustor liner facing the combustion gases is characterized by not having a film-cooling boundary layer of air to reduce quenching of the combustion gases for reducing exhaust emissions.
In the Scott and Vogt combustors, spent cooling air after impinging upon the combustor liner is discharged downstream as film cooling. In an integrated high performance turbine engine application, it would be advantageous to use substantially all the available air in the combustor for burning and thereby improve combustion efficiency while reaching stoichiometric combustion temperatures. Stoichiometric temperature refers to the maximum achievable gas temperature. The more available air to the combustor that is used for mixing with the fuel, the better the combustor efficiency. Conversely, as more available air to the combustor is used for cooling, the mere difficult it is to achieve stoiciometric temperature. In designing high thrust to weight ratio engines, higher combustor temperatures generally translate into higher thrust improvement for the same size engine.
It would therefore be desirable to provide a combustor cooling system in which cooling air after impinging upon an imperforate liner is transferred upstream to the combustion dome area for combining with fuel for burning.